Spacecraft Optimized Arc Rocket

FIG. 1 is a cut-away side view of a thruster according to the present invention.

FIG. 2 is a schematic view of a thruster in the configuration as a Hall current accelerator.

FIG. 3 is a schematic view of a thruster with associated support systems, some of which are subsystems of the carrier space vehicle.

FIGS. 4, 5 and 6 show three alternative constructions for a portion of the propellant flow path.

FIGS. 7 and 8 show two alternative embodiments of circuitry for the Hall current accelerator.

FIG. 9 shows an alternative embodiment of the thruster of FIG. 1 including the provision of an external preheater.

Referring to FIG. 1, a thruster assembly 10 according to the present invention is comprised of an anode 12, cathode 14 and a containment chamber 16. Attached to the containment chamber 16 is a propellant injector 18 through which propellant is conducted at ingress opening 19. The proplellent enters into the vortex shaped flow dividers and channels 20 causing a swirl or vortex flow pattern as well as a separation of flow into components. A portion of the flow at 22 swirls around the anode, bypassing the arc, regeneratively cooling the containment chamber wall through contact therewith. Another portion swirls to the base of the device and then enters the arc core flow channel 24 through bypass passages 26. The third portion of the flow is caused to enter the anode 12 through annular passages 28 to regeneratively remove anode heat through contact with the anode interior. Is this flow exits the anode 12 it enters the arc core flow channel 24, passes through the arc 30 and then enters the flow mixing-expansion section 34. This division of flow permits operation of the arc at lower voltage than would be required if all of the flow passed through the arc 30. This is because the lower flow rate caused by this division of flow has less tendancy to extend the arc path. Less mass going through and being heated directly by the arc column results in lower column voltage. The partially ionized and dissociated flow from the arc core flow channel 24 is combined with the flow that bypassed the arc area 30. This combining of gases in a mixing chamber will tend to equilibrate their temperatures and will efficiently convert most of the ionization and dissociation energy of the arc column gas to thermal energy in the bypass gas which is then converted to useful thrust in the nozzle.

Most of the divided flow is mixed in the flow mixing-expansion section 34; however, some of the cooler bypass flow will remain nearest the wall up through the nozzle 36 this maintaining a beneficial thermal gradient. The flow passes through the nozzle throat 38 and then out through the expansion nozzle 36. Referring still to FIG. 1, the flow divider and flow channels 20 may be formed as part of the containment chamber 16 and/or the anode 12 and/or include separate components which could also function as internal radiation shields. (In FIGS. 4-6, like elements are given like primed reference numerals) FIG. 4 shows a fragmentary view of the thruster showing containment walls 16'. The flow dividers and channels 20' including one outlet leading to location 22' and a second outlet leading to arc core flow channel 24' are shown formed as a part of the containment chamber 16'. Referring to FIG. 5, in this embodiment, the flow dividers and channels 20" are shown as formed as an integral part of the tubular anode 12". This embodiment shows only a single inlet and a single outlet for propellant flow leading to the arc core flow channel 24". Referring now to FIG. 6, this embodiment shows the flow dividers and channels 20'" formed as an integral part of the tubular anode 12'" and further including a first outlet for propellant adjacent the area 22'" bypassing the arc and a second outlet admitting propellant flow into the arc core flow channel 24'". The propellant is caused to have heat transferred to it as it flows through the device for any or all of the following purposes: to vaporize, decompose react or be heated to higher temperature. The flow channel surfaces may be coated or plated for the purpose of chemically isolating the passages 20 or walls 16 from the propellant or to enhance the decomposition or reaction of the fluid.

The function of the internal shields 40, external shields 42, and electrode shields 44 is to contain and focus radiated energy toward the hot core of the thruster 24, 30, 34. The surface of these shields will have a high reflectivity (low emissivity, .epsilon.<0.2).

The support structure 46 is attached to the containment chamber 16. FIG. 1, illustrates a possible version of a method to extend the length of this support. In this illustration, the support is two concentric tubes 48 and 50 connected at one end with a ring 52. These support tubes have material cutouts (not shown) to reduce thermal conduction. This structure may also be used to provide secondary support to the nozzle 36 by means of a low thermal conductivity support ring 54. The support ring 54 may, if desired, be made of an insulating material, such as, for example, boron nitride or a low thermal conductivity, light weight metal such as titanium. Thus, direct heat conduction through the supporting structure must take a long and tortuous path from the internal hot core 24, 30, 34 to the space vehicle mounting in interface 56. This interface will then be relatively cool, thus minimizing the heat conduction to the space vehicle via the device mount 46. An electrically insulative disc 55 extends within chamber 16 adjacent attachment 56 and structurally locates anode 12 and cathode 14 within the thruster 10.

In order to further reduce the temperatures to which the device mount 46 is exposed, the exterior surface of the containment chamber 16 is coated or surface conditioned to obtain a high emissivity coating. A preferred high emissivity coating would be hafnium carbide (HfC). The use of the high emissivity coating on exterior surface would reduce the radiative heat loss to the device mount 56. The reduction in temperature of the outer portion of the concentric support tube would also reduce the amount of radiation transferred away from the thruster to other portions of the space vehicle or to other external attachments or sensors (not shown).

By providing the high emissivity coating on exterior surface 16 and/or on the outer surface of the outer portion of the support tube 46, the operating temperature of the thruster 10 at the propellant hot core regions 24, 30, 34 can be increased even though maintaining a low temperature at exterior surface. This will also enable the thruster 10 to operate at higher efficiencies. It should further be noted that without the use of the folded arrangement of the concentric support tube 46 separated by the laminations of shields 42, it would be necessary to reduce heat loss at the exterior surface of the containment chamber 16 and a low emissivity coating on the exterior surface of the support structure 46 would be less practical. In addition to the support structure 46, various controls are attached to thruster 10. The reduction of temperature accomplished by the use of the high emissivity coating at exterior surface 46 reduces the maximum temperature to which external components of these controls are exposed.

When propellant is being expelled from the mixing and expansion chamber 34 to produce thrust, high temperatures are created at the nozzle throat 38. To withstand these high temperatures, it is necessary to use high temperature or refractory materials at that location. An option is to use a separate insert 38 for the nozzle throat area. The insert extends to connect the containment chamber 16 with the expansion nozzle structure 36. By using a separate insert 38, costly materials are only required for the hottest portions of the containment chamber 16. The expansion section of the nozzle 36 is exposed to a lower temperature. This part of the nozzle can be made as a separate section or shell which covers laminations of external radiation shields 42 which are located in that area. It can be seen that, because of the uncoupled heat-exchange relationship of the insert 38 and the expansion portion 36, the amount of heat loss through the expansion nozzle due to conduction and radiation during the operation of thruster 10 is reduced. In the preferred embodiment the insert 38 is made of thoriated tungsten. Further, the preferred material for the walls of chamber 16 and the nozzle inlet 17 is molybdenum/rhenium, and the expansion nozzle 36 is preferably made of TZM (a moderate cost molybdenum alloy) or titanium.

By separating insert 38 from the expansion portion 36, a means is provided to reduce radiation losses from the nozzle 36. The insert 38 is not mechanically joined directly to the expansion portion 36 and a blocking effect is accomplished by a thin diaphragm 58 between insert 38 and the expansion portion 36. As shown, the diaphragm 58 is located on the expansion side of the nozzle throat 36 and acts as a "block" to prevent propellant flow from entering the radiation shield area 42 through the gap that would otherwise be present between insert 38 and expansion portion 36. Thus, the diaphragm 58 acts as a layer of metal blocking the gap from flow-through. In the preferred embodiment, this diaphragm 58 will be made of a tungsten foil. Therefore less power is transferred from the hot insert 38 to the expansion portion 36.

It is expected that temperatures at the insert 38 will reach a range of 1700.degree.-2000.degree. K, whereas temperatures at the expansion portion 36 will reach a range of 1100.degree.-1400.degree. K. Without a separation of the nozzle portions 36, 38 energy losses would be represented by:

where

P=power-energy/unit time;

.sigma.=(Stefan-Boltzman) constant;

.epsilon..sub.n =integrated emissivity of the nozzle 36;

A.sub.n =effective area of the nozzle 36;

T.sub.n =temperature of the nozzle 36.

With the separated structure of the preferred embodiment, energy losses would be represented by:

where .epsilon..sub.i =emissivity 16 of the insert 38;

A.sub.i =area of the insert 38;

T.sub.i =temperature at the insert 38;

T.sub.P =temperature at the expansion portion 36.

These equations are approximate models because of such factors as thermal conductivity and direction of thermal radiation. Using an 0.1 pound thrust arc engine for an illustrative example, the typical power loss values for an integral nozzle without a diaphragm would be: ##EQU1## For a nozzle with a diaphragm, approximate values would be: ##EQU2## These examples indicate that the radiative power loss from the nozzle can be reduced by more than a factor of 2 using the diaphragm and thermal uncoupling.

To configure the arc thruster with a Hall Current accelerator extension, several minor changes are made to the basic thruster and several additional components are added to the device. These are illustrated in FIG. 2, wherein like elements are given like reference numerals with a prime numberal "4".

The current flow is initiated by thermionic emission from the conical tip of the cathode 14.sup.4. The cathode is surrounded by an anode which now functions as a buffer electrode 60. Gas is injected into this subassembly through the tube 62 and flows tangentially in the cavity 24.sup.4 and out of the current density control orifice 64. The cathode 14.sup.4 and buffer electrode 60 are mounted on an insulator 66 using "feed-thru's" that are sealed against leaks from internal pressure. This subassembly 60, 14.sub.4, 66 and 24.sup.4 is surrounded with electrical and thermal shields (isolator) 68. With gas injected thru the tube 62, the subassembly consisting of 60, 14.sup.4, 66, 24.sup.4, 64 abd 68 is an alternate arc-jet configuration and can be operated as such. A solenoidal magnet coil 70 mounted around the shields can be operated in conjunction with the arc-jet to help stabilize operation of the arc and spacecraft power supply combination.

With the inclusion of components 72, 74, 76 and 78 as shown in FIG. 2, the device can be operated in the Hall Current Accelerator mode. In this mode additional propellant flow is injected into the arc heated flow through inlet 76. The arc discharge is transfered from the buffer anode 60 to the accelerator anode 74. Magnets 70 and 78 are energized to create a common polarity, extended magnetic field. Magnets 70 and 78 can be excited by placing them in series with the arc current as shown in FIG. 7 or they can be separately excited as shown in FIG. 8. Switches 81 and 83 may be included if desired. The anodes 60 and 74 are connected in parallel with each circuit including power switches 73, 75 and/or power-limiting resistors well known in the art but are not operated simultaneously. The power supply is referred to by reference numeral 80. Referring now to FIG. 8, this figure shows the configuration with separate power supplies. As shown, a first power supply 82 supplies power to the magnet 78. A second power supply 84 is connected to the accelerator anode 74 and the cathode 14.sup.4. A third power supply 86 supplies power to the magnet 70 while a fourth power supply 88 is connected between the cathode 14.sup.4 and the buffer electrode 60. The configuration will operate as a Hall current accelerator. A portion of the propellant is injected into the buffer anode/ electrode 60 and is heated by regeneratively cooling the electrode 60 and then passing thru the arc area which is maintained between the cathode tip and inner diameter of the current control orifice 21. This mode of operation can also be used to pre-heat the accelerator anode 74 if the device is also used for thermal decomposition of an exothermic propellant (e.g. hydrazine). The electrical connections shown in FIG. 7 and 8 may be so operated in conjunction with propellant inlet valves(not shown) in a pulsed mode as described in the "Summary of the Invention."

FIGS. 1, 2 and 3 are used to illustrate the operating procedure of the arc jet portion of the invention. Prior to operating the augmentation arc 30, propellant from a space vehicle storage tank 90 must enter and flow through thruster 10. In the case where the propellant must be vaporized, decomposed and/or reacted, this may be accomplished in an external chamber 92, or within the injector assembly at the intersection with flow divider section 20. If this vaporization, decomposition and reaction is induced by thermal energy, an external heat source 94 would be provided by the space vehicle's solar cells or batteries. Such an initial phase of warm-up may typically last several minutes. If catalytic decomposition is used, no prior warm-up is required. With propellant flow in thruster 10, the arc 30 can be initiated. This arc initiation can be accomplished by means of a high-voltage single discharge, or by a high-voltage, high-frequency discharge from an external source 96 coupled to the power leads 98 and 100.

Following arc initiation 30 the arc will be maintained by the space vehicle power supply 102. This power supply 102 may include a power controller and/or regulator (not shown) in addition to the energy supply and storage system, likely batteries. Operating parameters of thruster 10 must be matched to this supply 102 with a minimum of power conditioning or regulation for maximizing overall power efficiency. Specific features of thruster 10 have been included to optimize interaction with standard space vehicle power supplies 102 requiring none or minimum conditioning. These include the variable arc gap and the division of propellant flow.

Typical space vehicle propellant supply systems 90 are of a "blow-down" design; that is propellant supply pressure, and hence the flow rate, will decrease over mission lifetime. The arc augmentation thruster 10 has sufficient heat capacity and a self-adjusting arc gap 30 which operates as is well known the art to accommodate a wide range of "blow-down" conditions.

When the device is operating in the Hall Accelerator mode, a "virtual cathode" 104 forms which consists of ionized gas flowing along and expanding radially across the magnetic field lines. The expansion rate is such that the combined radial forces due to the self-magnetic field and the azimuthal Hall currents in the jet are adequate to radially confine the pressure exerted by the gas in the cathode jet. Electromagnetic "pumping" forces in the cathode jet accelerate the ions axially against the electric field. Elastic and ionizing collisions between atoms, ions and electrons in the cathode jet also permit radial expansion of the cathode jet; hence the rate of radial expansion can be controlled by adjusting the mass flow rate of propellant injected through the tube 62. Generally speaking, about 1/1Oth of the total flow will be injected through the inlet 62 and 9/10th's will be injected through the inlet 76. With this general ratio being followed, stable operation of the device will result. Other ratios may be used, if desired.

Once the cathode jet reaches region 106, the axial electric field goes through a "zero" and the electron flow through the cathode jet divides; part flows with ions from the anode sheath 108 and from the cathode jet into the neutral exhaust beam 110, the remainder enter the anode sheath 108 and flow toward the anode 74, continuing to gain energy. As they enter the anode housing 72, 74 in the form of an annular sheath they encounter many atoms. Because of the high electron energy, most electron-atom collisions are ionizing collisions. The ions thus produced are accelerated radially by the strong radial electric field in this region, but are turned azimuthally by the solenoidal magnetic field. The expanding solenoidal magnetic field and the positive axial electric field in the anode sheath both act to accelerate the ions axially away from the anode. These ions eventually enter the plasma exhaust beam. The anode housing 74 is made long enough to confine the propellant injected through 76 long enough to have most of it ionized by passing through the anode sheath. The distance between the cathode-buffer (21) and the position where the current terminates (31) is given approximately by: ##EQU3## Values for the above quantities could be: ##EQU4## If the magnetic field were increased from 1000 gauss to 2000 gauss, the distance L would increase to 6.63 inches.

The arc current, the mass flow rate and the strength of the applied magnetic field are adJusted so that they meet the requirements of the minimum voltage hypothesis for single ionization of the propellant atoms. The plasma beam then has an exhaust velocity equal to the Alfven speed of the gas. While conducting experiments with an arc transverse to a magnetic field, Alfven found that he could not increase the velocity of the discharge through the gas beyond a certain value. This "critical" speed was different for different gases, but correlated using the formula: ##EQU5## where, .vertline.e.vertline.=charge on electron.

V.sub.I =first ionization potential of the atom

m.sub.I =mass of the ion.

This critical velocity, v.sub.cr, became known as the Alfven speed. When the implications of the minimum voltage hypothesis were worked out for plasma accelerators by applicant, the analysis indicated that the exhaust velocity of the plasma beam out of the accelerator should be equal to this critical or Alfven velocity, provided adequate propellant mass flow was available. If inadequate mass flow were provided, then a number of things could happen:

(1) The electron temperature could rise and ionize in multiple fashion the propellant atoms (resulting in a higher effective Alfven velocity).

(2) The discharge could increase in volume and ionize ambient gas which was then injected into the plasma beam.

(3) The discharge could evaporate and ionize material from the cathode, or the anode, or from insulators and inject this into the plasma beam.

(4) The discharge could go unstable.

FIGS. 1-3 can be used to illustrate the test procedure of the invention. Inasmuch as thruster 10, according to the present invention, is designed to operate in a vacuum environment, a special test facility is provided for ground testing. Thruster 10 is placed in a vacuum enclosure (not shown) with an outside exhaust duct (not shown) to provide connection with vacuum source (not shown). An alternate configuration (not shown) would duct the propellant exhaust directly from the expansion nozzle to the external vacuum supply. This method would permit achieving better vacuum conditions within vacuum enclosure by use of typical low flow, hard vacuum pumping systems such as by absorption pumps and/or ion pumps. Unlike augmented devices heated by free-standing coil heaters, ground testing in a terrestrial gravity environment will not impose any additional test or facility constraints.

What has been described is a preferred embodiment of the invention. It should be noted that it is possible to provide various other arrangements. For example, if the fuel is injected into thruster 10 in the unreacted liquid state, an external heater 120 would be used as shown in FIG. 9 to pre-heat thruster 10 to a safe temperature above the thermal decomposition temperature, about 1000.degree. K. This internal coupling of exothermic decomposition and electrical performance augmentation eliminates the heat losses from an external mounted decomposition chamber and from the connecting injection tube.

The foregoing detailed description is to be clearly understood as being given by way of illustration and example only, and the spirit and scope of this invention is intended only to be limited by the scope of the following appended claims.